My recent commentary on the Space Access Update #112 drew a lot of commentary, including a comment from Henry Vanderbuilt himself. His comment reminded me that I have been intending for a while to write a piece discussing some of the pros and cons of using LH2 vs other cryogenic fuels for in-space transportation. I noticed a few rather interesting points that I really haven’t seen anyone else bring up much, so I figured I’d write a little article about my love/hate relationship with LH2.
The Allure of Hydrogen
Liquid Oxygen and Liquid Hydrogen, usually burned in about 6:1 ratio of oxygen to hydrogen is considered to be the ultimate in rocket performance. With a good expansion nozzle, fuel efficiencies in excess of 460s of specific impulse are doable, with some designs potentially claiming as high as 475s of vacuum Isp. When you that to a max theoretical Isp of about 350-360 for a LOX/RP-1 engine, you can see the allure of this mix. NASA in particular has been very fond of this mixture. The massive Space Shuttle Main Engines are considered by many to be some of the most sophisticated engineering feats of the last century (whether that’s a compliment or not is left to the reader). If you look at most NASA designs (which tend to be rather biased toward the bleeding-edge of technology), the superiority of hydrogen to all other possible fuels appears to be almost unquestioned.
However, starting in the early 90s, this orthodoxy began to be questioned. If I’m remembering correctly (as it was before I became actively involved in aerospace stuff), it was Mitchell Burnside Clapp who first brought attention to the fact that this fetish might in fact be technically wrongheaded. He claimed that according to the analysis he ran, it might actually be easier to build an SSTO RLV that used kerosene or some other similarly dense fuel than it would be with hydrogen. Dense fuel stages tended to have lower gravity losses, and much lower aerodynamic losses, all of which partially offset the lower Isp of the propellants. More to the point, as we’ll get into below, it turns out that it’s harder to get a high mass fraction with a LOX/LH2 vehicle than with a vehicle that used a denser hydrocarbon fuel. [Ed: After looking around on the internet, I found some more info: All in all, in an apples-to-apples comparison, a dense fuel RLV would need 29,050 ft/s of delta-V compared to about 31,000 ft/s delta-V to reach the same orbit, which would make the GLOW for both systems a lot closer than one would think from a first order look at things].
Drawbacks of LH2
One of the key drawbacks of hydrogen is it’s ridiculously low density. Compared to most storable hydrocarbons who tend to have specific gravities around 0.7-0.8, hydrogen’s specific gravity is a measly 0.07! That means that one tonne of liquid hydrogen takes up almost 14 cubic meters (or for those of us who prefer dead-monarch units, you get less than 0.5lb of the stuff per gallon). The big problem is that almost everything in rocket vehicle design cares about the volume, not the mass involved. Tanks mass scales almost linearly with volume. Pumps pump volume, not mass. Feedlines have to be sized for the volumetric flow rate of the fluid. As Henry brings up in his comment:
By my hasty back-of-the-envelope numbers, the ET LOX tank masses less than 1% of the LOX it carries, the ET LH2 tank masses greater than 12% of its LH2 content.
Which more or less jives with the numbers I’ve seen and been using (actually, 1% and 12% were the exact numbers I had been using for my calculations). Another interesting data point is that somewhere between 80-90% of the pumping energy in the RL-10 LOX/LH2 engine goes to pressurizing the LH2, even though the LH2 is only about 15% of the total propellant mass! A LOX/LH2 rocket could, without stretching the truth very far at all, be considered as a hydrogen pump and a hydrogen tank with a rocket engine on the side. Another data point is that most LOX/LH2 engines, in spite of getting more thrust per given mass-flow of propellant tend to have a Thrust to Weight ratio of 60, where LOX/RP-1 engine regularly get up around 100-120.
There’s another annoying problem with LH2–the stuff is so darn cold. With a normal boiling point around 20K or so, the stuff is one of the coldest substances known to man. Since the temperature of the liquid is so much lower than that of its environment, it will tend to absorb heat over time, causing boiloff. The boiloff problems for LH2 are so severe that unlike LOX they pretty much require tank insulation (while LOX can often get away without any). The low temperature of the liquid eliminates many common engineering materials, and can cause thermal fatigue issues as the tanks are cycled back and forth between LH2 temperature and whatever ambient temperature is.
Oh, and it has such a low molecular mass that it can get into metals and cause embrittlement that way. Oh, and it makes sealing tougher. Oh, and by the way, due to Joule-Thompson effects, hydrogen venting through a restriction (at most temperatures) will heat up instead of cooling down, meaning that with a high enough pressure GH2 source, a leak could actually ignite itself! Oh, and it burns with a nearly invisible flame that is several thousand K…
There are probably more problems with Hydrogen, but I think I’ve already brought up some of the worst.
So What are the Alternatives?
Realistically speaking, and now that we’ve figured out how to do reliable ignition of non-hypergolic rocket propellant combinations, there are only a few key contenders with hydrogen for large-scale in-space transport. Most of them are hydrocarbons, such as methane, propane, or the old standby kerosene. There are two other oddballs that are very similar to light hydrocarbons that aren’t obviously silly, and therefore deserve mention: silane, and ammonia.
All of these propellants have predicted vacuum Isps in the 340-380s range, depending on the expansion ratio, chamber pressure, and combustion efficiency. All of them have bulk propellant densities much better than LOX/LH2. Ranging from a bulk density of about 1.03 for LOK/RP-1, down to 0.83 or so for LOX/Methane, as compared to 0.33 or so for LOX/LH2. That means you can get somewhere near 2.5-3x as much propellant into the same volume when compared to LH2. This is important for two things: drylaunch, and tank mass.
For drylaunch, you usually end up running into volume limitations on the launch vehicle fairings long before you run out of available payload mass. For example, the Atlas V, 4.5m PLF has about 180 cubic meters of space in its cylindrical section. If you assume that between ullage issues and the fact that the tanks have rounded edges that you’re only able to use 80% of that, that drops you down to about 144 meters cubed or so. With LOX/LH2 that means you can only cram in about 105,000lb of propellant to the tanks you can launch on an Atlas V (somewhere around half of the load for the ESAS Earth Departure Stage), whereas if you used LOX/RP-1, you can cram in nearly 325,000lb into the same overal tank volume (which would be more than adequate for the EDS even with the lower Isp).
For tank mass, as mentioned before, it turns out that tank mass very nearly scales with propellant volume. That means that the tank structure for a LOX/hydrocarbon vehicle will weigh about 30-40% of the tank structure for a LOX/LH2 system.
Another important thing is boiloff. Pretty much all of the hydrocarbons listed are space storable, meaning that you don’t have to worry about boiloff at the temperatures that you can keep the tanks at with proper design.
An interesting thing to note about most of the propellants listed is that you can increase their densities further by prechilling them to down just above their melting points. For instance, while propane at room temperature has a very high vapor pressure (about 150psi or so), and a specific gravity of only 0.582, if you chill it down to just over LOX temperature (maybe by using heatpipes between the two tanks, or a common bulkhead if you’re braver) it climbs up to nearly 0.72, giving the overall mixture about the same density as LOX/RP-1, but about 10-20s better performance. [Ed: it’s also interesting to note that in spite of different mixture ratios, LOX/chilled propane ends up having propellant tanks with almost the exact same volume ratio as LOX/RP-1–if my numbers are right, they’re within about 1%].
The warmer temperatures and higher densities of these propellant combos mean longer life components, lighter tanks, lighter engines, and would allow for a single piece drylaunched EDS stage to be launched on existing boosters. Not to mention cheaper to design, easier to handle, etc.
Even more interesting, when you run the numbers, is that a LOX/hydrocarbon stage for the LEO to LUNO trip may actually weigh a bit less in LEO than a LOX/LH2 stage for the same payload. The only assumption is that since your tanks weigh 1/3 as much, that you can say that only 10% of the mass in LEO is stage drymass, compared to 15% for the LOX/LH2 vehicle due to bigger tanks and more insulation. Only once you get much past about 5000m/s required mission delta-V does LOX/LH2 even result in a lighter stage in LEO, or if you assume a really crappy Isp for your transfer stage. [Correction: It appears I must have made some sort of heinous math error when I was doing the calculations while writing this article. Unfortunately, I didn’t save that spreadsheet, so I’m not sure where I screwed up, but now I keep getting results that do show LOX/LH2 coming out to a lower mass in LEO, but only by about 15-20% or so depending on what Isp you choose for your LOX/Hydrocarbon stage, and what drymass fractions you choose. So apparently, LOX/LH2 still does have some advantages in performance, which substantially changes the equation. Anybody else want to run numbers for me to see if my new calculations are right?]
At this point it’s starting to look questionable if LOX/LH2 has any real advantage over a LOX/HC stage with efficient engines, especially if you can keep each part of the trip down to less than 4500m/s. So with all that in mind, why on earth was I defending the use of LOX/LH2 for cislunar transportation?
LH2: What’s there to Love?
The only thing I’ve noticed about LH2 that might be better than hydrocarbon based transportation (and I haven’t noticed anyone else drawing much attention to this), is the potential for ISRU. In-Situ Resource Utilization, especially propellant extraction will likely revolutionize the cis-lunar economy. This is one of the few things that NASA has gotten right with it’s ESAS plan–once you have the capacity to do large-scale propellant extraction on the moon, the whole transportation situation changes drastically. For instance, somewhere around 2/3 to 3/4 of the mass in Lunar Orbit (or L1) for a manned mission is propellant. Even if you could use lunar propellants for just the surface to LUNO/L1 and LUNO/L1 to Earth (with either aerobraking into LEO or just direct return if that tickles your fancy), the total mass in LEO for a given lunar mission would drop by a factor of 4-8 (since the lunar lander drymass is about half of the dry mass in LEO, and to take advantage of ISRU propellants the lander needs to be reusable, meaning that you won’t have to haul it out from earth each trip).
There’s one big problem. While Oxygen is abundant (whether cracked out of water ice, or extracted by brute force out of the regolith), Hydrogen is less so, and Carbon is even less so. Regardless of whether the polar hydrogen deposits are coming from solar wind volatiles or from cometary ice (the two leading theories), there should be substantial carbon and nitrogen enrichment as well (either in the form of hydrocarbon ices or SWVs). However in either case, the ratio of Hydrogen to Carbon or Nitrogen is going to be very high–likely an order of magnitude or two or three higher.
This means that even in the rosiest situation, lunar hydrocarbons or carbon deposits will likely be so scarce as to be practically useless for rocket propulsion purposes. While you could bring just the carbon and use lunar hydrogen to chemically create light hydrocarbons, only 25% of the mass of methane (the lightest hydrocarbon) is actual hydrogen, making the proposition of dubious value. Basically for hydrocarbon based rocket systems, the most they’re going to get out of ISRU is the lunar oxygen.
And that is the second problem. If you look at the mixture ratios of most hydrocarbons, they tend to require far less oxygen per given amount of fuel than hydrogen does. For LOX/LH2, the ratio is usually 6:1, whereas for LOX/Methane it is only 3.4:1, 3.1:1 for LOX/propane, and only 2.7:1 for LOX/RP-1. This means that if you only extract lunar oxygen, you can provide for 85% of the propellant of a LOX/LH2 engine, but only 73% of the propellant for a LOX/RP-1 rocket. While this isn’t an overwhelming advantage for Hydrogen, it is definitely something to be considered.
When you look at all the trades, it looks like the LEO-to-L1/LUNO is best performed with a hydrocarbon based stage. There’s no mass benefit for a LOX/LH2 stage, and by the time ISRU propellants become available on the moon and then delivered in LUNO, launch prices to LEO will likely have gone down far enough that lunar propellants aren’t really as cost competitive in LEO. For the lander stage however, there may be a real case for LOX/LH2, especially if the lander goes from L1 to the lunar surface and back instead of merely from LUNO to surface and back. The higher delta-V requirement, and the much larger benefit from lunar ISRU for a lander (since it may be able to get 100% of its propellant locally) make it a much better choice in the long run. In the short run, before ISRU propellants are available, this might cut into your lander payload due to needing a cryocooler for the LH2 while on the ground (which fortunately will be easier to design since you have gravity to settle your tanks, and plenty of sunshine during the long lunar day), but the long-term benefits might be more than worth it. Ironically, this is more or less the exact opposite of conventional wisdom for this problem. [Ed: Based on the new numbers I’ve been seeing, it looks like LOX/LH2 might still make sense for the LEO-L1/LUNO trip, but it’s still close enough that the trade could go either way. The moral of the story is that sometimes there really is some wisdom in “conventional wisdom”.]
Thoughts, comments, flames?
Latest posts by Jonathan Goff (see all)
- Random Thoughts: A Joint International Debris Remediation Effort - October 22, 2021
- Unorthodox Reusable Lunar Landers Concepts - June 12, 2021
- Goff Family 2021 Summer Sabbatical Part 1: Utah Trip - June 1, 2021